U.S. patent number 9,347,400 [Application Number 13/684,729] was granted by the patent office on 2016-05-24 for rocket stage having engine thrust frame integrated with tank.
This patent grant is currently assigned to Astrium GmbH. The grantee listed for this patent is Astrium GmbH. Invention is credited to Jesus Gomez Garcia, Markus Jaeger, Menko Wisse.
United States Patent |
9,347,400 |
Jaeger , et al. |
May 24, 2016 |
Rocket stage having engine thrust frame integrated with tank
Abstract
A rocket stage for a spacecraft includes an engine, a tank for
storing a fuel and an oxidizer for combustion in the engine, a
rocket stage primary structure, and an engine thrust frame that
connects and transmits forces between the engine and the primary
structure. The engine thrust frame includes at least an internal
part thereof arranged internally within the tank. This internal
part of the engine thrust frame forms an imperforate partition that
divides an interior space of the tank into at least two chambers
for storing the fuel and the oxidizer separately from one another.
Propellant management devices including liquid guide vanes and
refillable liquid reservoirs may be provided in connection with the
partition in the tank interior space.
Inventors: |
Jaeger; Markus (Bremen,
DE), Wisse; Menko (Bremen, DE), Garcia;
Jesus Gomez (Bremen, DE) |
Applicant: |
Name |
City |
State |
Country |
Type |
Astrium GmbH |
Taufkirchen |
N/A |
DE |
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Assignee: |
Astrium GmbH (Taufkirchen,
DE)
|
Family
ID: |
46845573 |
Appl.
No.: |
13/684,729 |
Filed: |
November 26, 2012 |
Prior Publication Data
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Document
Identifier |
Publication Date |
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US 20130263573 A1 |
Oct 10, 2013 |
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Foreign Application Priority Data
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Nov 25, 2011 [DE] |
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10 2011 119 921 |
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Current U.S.
Class: |
1/1 |
Current CPC
Class: |
F02K
9/60 (20130101); F02K 9/42 (20130101); F02K
9/605 (20130101) |
Current International
Class: |
F02K
9/60 (20060101); F02K 9/42 (20060101) |
References Cited
[Referenced By]
U.S. Patent Documents
Foreign Patent Documents
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1270304 |
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Oct 2000 |
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CN |
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1421673 |
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Jun 2003 |
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CN |
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101268302 |
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Sep 2008 |
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CN |
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WO 2010/037561 |
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Apr 2010 |
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WO |
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Other References
Chinese Search Report issued May 19, 2015 in Chinese Patent
Application 2012104756343. cited by applicant.
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Primary Examiner: Meade; Lorne
Attorney, Agent or Firm: Fasse; W. F.
Claims
What is claimed is:
1. A rocket stage for a spacecraft, comprising: an engine; a tank
for storing a fuel and an oxidizer for combustion in said engine; a
first refillable liquid reservoir; a second refillable liquid
reservoir; a rocket stage primary structure; and an engine thrust
frame that connects, and transmits forces between, said engine and
said primary structure; wherein said engine thrust frame includes
at least an internal part thereof arranged internally within said
tank, wherein said internal part forms an imperforate partition
that divides an interior space of said tank into at least two
chambers for storing said fuel and said oxidizer separately from
one another, wherein said first refillable liquid reservoir is
arranged in communication with a first one of said at least two
chambers and said second refillable liquid reservoir is arranged in
communication with a second one of said at least two chambers, and
wherein said second refillable liquid reservoir is arranged
concentrically with said first refillable liquid reservoir at an
end of said tank adjoining said engine and, said first refillable
liquid reservoir and said second refillable liquid reservoir are
bounded and separated from one another by said internal part of
said engine thrust frame.
2. The rocket stage according to claim 1, wherein said primary
structure is arranged at an opposite end of said tank relative to
said end of said tank adjoining said engine.
3. The rocket stage according to claim 1, wherein no part of said
engine thrust frame extends externally around said tank from said
engine to said primary structure for transmitting forces between
said engine and said primary structure.
4. The rocket stage according to claim 1, wherein said internal
part is an entirety of said engine thrust frame, and all of said
engine thrust frame is arranged internally within said tank.
5. The rocket stage according to claim 1, wherein said engine
thrust frame further includes a first external part thereof
arranged externally outside of said tank, and wherein said first
external part is configured as a conical ring and connects said
internal part for force transmission to said primary structure.
6. The rocket stage according to claim 1, wherein said engine
thrust frame further includes a second external part thereof
arranged externally outside of said tank, and wherein said second
external part is configured as a conical ring or a cone and
connects said internal part for force transmission to said
engine.
7. The rocket stage according to claim 1, wherein said engine
thrust frame further includes an external part thereof arranged
externally outside of said tank, and wherein said external part is
connected directly to or extends continuous with said internal part
whereby said engine thrust frame penetrates through a tank wall of
said tank.
8. The rocket stage according to claim 1, wherein said engine
thrust frame further includes an external part thereof arranged
externally outside of said tank, and wherein said external part is
discontinuous from said internal part with a tank wall of said tank
interposed therebetween, whereby said external part and said
internal part are aligned with one another and connected on
opposite sides of said tank wall for force transmission
therebetween through a thickness of said tank wall.
9. The rocket stage according to claim 1, wherein said primary
structure has a cylindrical configuration and is connected to said
internal part of said engine thrust frame or to said tank adjoining
said internal part for force transmission between said internal
part and said primary structure, without a conical attachment ring
interposed between said primary structure and said internal part or
said tank.
10. The rocket stage according to claim 1, further comprising
propellant guide plates that are adapted for collecting and guiding
the fuel or the oxidizer, wherein the propellant guide plates
protrude inwardly from and extend along said internal part of said
engine thrust frame in said first one of said chambers.
11. The rocket stage according to claim 10, further comprising
additional propellant guide plates that are adapted for collecting
and guiding the oxidizer or the fuel, wherein the additional
propellant guide plates protrude inwardly from and extend along a
tank wall of said tank in said second one of said chambers.
12. The rocket stage according to claim 1, further comprising a
disc-shaped reinforcement plate or rim interposed between and
connecting said engine thrust frame to said engine.
13. The rocket stage according to claim 1, wherein said tank has a
substantially spherical shape.
14. The rocket stage according to claim 1, wherein said tank
includes a cylindrical tank wall portion and two end domes at
opposite ends of said cylindrical tank wall portion.
15. The rocket stage according to claim 1, wherein said tank
includes two substantially elliptically curved domes.
16. The rocket stage according to claim 1, wherein said tank
includes two substantially partially spherically curved domes.
17. The rocket stage according to claim 1, wherein said internal
part of said engine thrust frame has a substantially conical shape
with a conical wall that extends linearly.
18. The rocket stage according to claim 1, wherein a vertex end of
said engine thrust frame ends inside said tank, and further
comprising a cylindrical extension that connects said vertex end to
said engine.
19. The rocket stage according to claim 1, wherein said internal
part of said engine thrust frame has a substantially conical shape
with a conical wall that extends convexly.
20. The rocket stage according to claim 1, wherein said internal
part of said engine thrust frame has a substantially conical shape
with a conical wall that extends concavely.
Description
PRIORITY CLAIM
This application is based on and claims the priority under 35 USC
119 of German Patent Application DE 10 2011 119 921.0, filed on 25
Nov. 2011, the entire disclosure of which is incorporated herein by
reference.
FIELD OF THE INVENTION
The invention relates to a rocket stage for operating spacecraft
with a primary structure, a tank for the separated storage of fuel
and oxidizer, an engine and an engine thrust frame that connects
the various components with one another.
BACKGROUND INFORMATION
The launch vehicle or booster rocket stages presently in use with
liquid propulsion systems include a series of components that
fulfill various different functions. This includes tanks for
storing fuel and oxidizer, whereby the fuel and the oxidizer are
stored either in separate tanks or in one tank with a partition
separating wall between fuel and oxidizer, as well as a so-called
engine thrust frame, designated as an Engine Thrust Frame (ETF) in
English. The latter transmits the loads of the engine into the tank
and into the primary structure on the one hand, and the loads of
the launcher or booster via the primary structure into the tank and
the engine on the other hand. Conventionally, the engine thrust
frame is separate from and external to the propellant tank(s), and
e.g. extends externally around the tank from the engine to the
primary structure so as to transmit forces from the engine to the
primary structure and vice versa. Furthermore, such rocket stages
include propellant handling or management devices, designated as
Propellant Management Devices (PMD) in English, which serve, among
other things, for ensuring the supply of gas-free and particle-free
propellant to the engine even when there is only a minimum
remaining propellant residual quantity, or for suppressing the
spinning of the liquid. It is desirable to reduce the size, mass
and complexity of the overall structure in order to reduce the
cost, increase the payload capacity and/or allow the shape and
configuration of the rocket stage to be optimized.
SUMMARY OF THE INVENTION
In view of the above it is an object of the invention to further
develop a rocket stage of the initially mentioned type, so that it
has a lowest possible weight and a compact design and causes the
lowest possible costs. Furthermore, the invention aims to avoid or
overcome the disadvantages of the prior art, and to achieve
additional advantages, as apparent from the present specification.
The attainment of these objects is, however, not a required
limitation of the invention or inventions defined in the appended
claims.
The above objects have been integrated in a rocket stage according
to the invention, wherein at least a part of the engine thrust
frame is arranged in the interior of the tank, whereby this
internal part of the engine thrust frame within the tank forms an
imperforate partition that is embodied completely closed or solid
in such a manner so that regions or chambers separated from one
another by the partition are formed in the tank for respectively
separately storing the components of the propellant. The term
imperforate means that the partition has no holes or openings
penetrating therethrough, through which a propellant component
(e.g. fuel or oxidizer) could flow, so that the chambers are
separated from one another by the imperforate partition.
By integrating the tank and the engine thrust frame into one unit,
the invention achieves an extremely compact arrangement and
simultaneously a considerable weight reduction. Moreover, in this
manner also the weight of the adjoining structures can be
correspondingly reduced. Through the compact design, the total
length of the inventive rocket stage is considerably smaller than
that of conventional configurations, which in turn effectively and
permanently reduces the mechanical loads such as wind loads and the
loads induced by the payload. Additionally, through the inventive
embodiment of the rocket stage, the various different functions are
optimally distributed to the individual components. Thus, the
substantially conically shaped tank-internal part of the engine
thrust frame is used not only for transmitting thrust forces
between the engine and the primary structure of the rocket stage,
but also as a partition in the tank for separating the fuel and the
oxidizer in separate chambers in the tank. Moreover it reduces the
geometric residuals.
Through the integration of also the propellant management devices
provided in an advantageous further development of the invention,
the compactness of the arrangement is still further improved.
Simultaneously, the components of the propellant management
devices, such as guide plates and refillable liquid reservoirs, are
also used for the force introduction. Namely, these propellant
management devices also supplement of contribute to the mechanical
strength and load-transmitting capacity of the internal part of the
engine thrust frame. Finally, the rocket stage according to the
invention has the reignition capability that is necessary in
special cases.
BRIEF DESCRIPTION OF THE DRAWINGS
In order that the invention may be clearly understood, it will now
be described in further detail, in connection with example
embodiments, with reference to the accompanying drawings,
wherein:
FIG. 1 shows a part of a rocket stage in a sectioned
illustration,
FIG. 2 is an enlarged illustration of a partial area of the
arrangement according to FIG. 1,
FIG. 3 is a different view of an enlarged partial area of the
arrangement according to FIG. 1, in the vicinity of the bottom
outlets of the tank, showing only a cutaway bottom portion of the
internal part of the engine thrust frame and of the outer tank
wall, with propellant guide plates arranged therein,
FIG. 4 is a further enlarged illustration of a partial area of the
arrangement according to FIG. 3,
FIGS. 5-7 are schematic sectional illustrations of various shapes
of tanks for a rocket stage,
FIGS. 8, 9 are schematic sectional illustrations of alternative
arrangements of engine thrust frames in tanks, and
FIGS. 10, 11 are two further schematic partial sectional
illustrations of alternative engine thrust frame shapes for a
rocket stage.
DETAILED DESCRIPTION OF PREFERRED EXAMPLE EMBODIMENTS AND OF THE
BEST MODE OF THE INVENTION
In the figures, the same components or components corresponding to
one another have the same reference numbers.
FIG. 1 shows, in a schematic illustration, a vertical section
through an inventive rocket stage, in which a substantially
conically shaped engine thrust frame 2 is joined onto an engine 4
and a primary structure 1, e.g. a main load carrying structure, of
the rocket stage. Thereby, the engine thrust frame 2 transmits the
arising loads or forces between the engine 4 and the primary
structure 1. A propellant tank 3 is spherically shaped in the
embodiment illustrated here, and stores propellant components, i.e.
fuel 6 and oxidizer 7 for combustion in the engine 4. A predominant
or main part, namely an internal part 2b of the engine thrust frame
2 is arranged in the tank interior space of the tank 3 and thereby
integrated in or with the tank 3. The engine thrust frame 2 in this
embodiment further includes an outer ring 2a, which is preferably
shaped as a conical frustum ring, and is located outside of the
tank 3 and connected to the primary structure 1 of the rocket
stage. The internal part 2b and the external part 2a of the engine
thrust frame 2 are connected with one another in a
load-transmitting or force-transmitting manner. This can be
achieved in different ways. For example, in one embodiment the
internal part 2b and the external part 2a are continuously directly
connected with one another, or continuously integral with one
another as one piece, whereby the engine thrust frame penetrates
through a tank wall of the tank 3, which must be suitably connected
and sealed in a pressure tight manner at the penetration junction.
As another example, in another embodiment the internal part 2b and
the external part 2a are separate and discontinuous from one
another but are aligned with one another and connected on opposite
sides of the tank wall of the tank 3, so that forces are
transmitted through the tank wall between the internal part 2b and
the external part 2a.
As can be seen from the figure, in this arrangement, the tank 3 and
engine thrust frame 2, which are typically arranged separately from
one another in conventional rocket stages, are here inventively
connected and integrated with one another, without being limited or
impaired in their respective functions. In fact, the inventive
integration of the engine thrust frame 2 into the tank 3 gives the
internal part 2b of the engine thrust frame 2 two additional
functions, as will be described in the following.
As a first additional function, the internal part 2b of the engine
thrust frame 2 serves as a partition or separating wall that
divides the interior space of the tank 3 into two regions or
chambers I and II for separately storing the propellant components.
For example, the first region or chamber I stores the oxidizer 7,
and the second region or chamber II stores the fuel 6. Thereby an
arrangement of several separate tanks or the installation of an
additional partition or separating wall in the single tank is not
necessary. Furthermore, by providing an intelligent pressure
control between the two tank regions I and II, the structural mass
of the conically shaped separating wall 2b (i.e. the internal part
2b of the engine thrust frame 2) can be reduced to a minimum.
In that regard, the substantially spherical shape of the tank 3
illustrated in FIG. 1 represents the optimal shape of a pressurized
tank for storing fuel and oxidizer, insofar as exclusively the
occurrence of pressure loads is being considered. Because the
engine 4, as also shown in FIG. 1, is connected with the rest of
the rocket stage only at one location (i.e. at the bottom vertex of
the substantially conical engine thrust frame), and the primary
structure 1 has an essentially cylindrical shape, therefore the
conical shape of the engine thrust frame 2 is the optimal shape for
transmitting forces between the cylinder of the primary structure 1
and the single connection point for the engine 4. Thus, by
arranging a major part 2b of the engine thrust frame 2 internally
within the tank 3, this allows the force transmission paths to be
optimized (e.g. with the conical shape of the internal part 2b of
the engine thrust frame, and the spherical shape of the tank wall
of the tank 3), and therefore allows the structural dimensions of
the engine thrust frame and the tank wall to be minimized for
carrying the arising loads.
Furthermore, as another additional function, the internal part 2b
of the engine thrust frame 2 participates in the propellant
management functions. Namely, the arrangement illustrated in FIG. 1
incorporates propellant management devices 5 including propellant
guide plates 5a and at least one refillable liquid reservoir 5b
connected or integrated with the internal part 2b of the thrust
frame 2. The guide plates 5a extend radially to the central axis of
symmetry of the rocket stage and connect the widest part of the
tank 3 with the tank outlet, in order to suppress the spinning of
the respective liquid propellant component and in order to collect
and guide it to the respective tank outlet. The refillable liquid
reservoirs 5b are secured above the tank outlets and hold a
sufficient liquid quantity at the outlet, on the one hand to ensure
that the liquid supplied from the tank is gas-free due to surface
tension forces, and on the other hand to ensure that propellant is
immediately available for a renewed ignition of the engine 4, even
before the engine's burn thrust causes an acceleration that drives
the rest of the propellant to the tank outlet. In that regard, the
guide plates 5a and the refillable liquid reservoirs 5b are
arranged in both regions I and II of the tank 3. Namely, guide
plates 5a1 are arranged on the thrust frame internal part 2b in the
chamber I for the oxidizer 7, and guide plates 5a2 are arranged on
the tank wall of the tank 3 and partially span like a web to the
thrust frame internal part 2b in the chamber II for the fuel 6. A
reservoir 5b1 for the oxidizer 7 is arranged at the bottom vertex
of the thrust frame internal part 2b, and a reservoir 5b2 for the
fuel 6 is arranged at the bottom of the tank wall of the tank 3
concentrically or coaxially around the reservoir 5b1. The several
propellant management devices 5, i.e. 5a1, 5a2, 5b1 and 5b2 provide
typical capillary and surface tension effects to accumulate, guide
and hold the fuel 6 and oxidizer 7 in the area of the tank chamber
outlets.
The enlarged illustration of a partial area of the above described
arrangement in FIG. 2 shows how possible residual quantities of
fuel 6 and oxidizer 7 accumulate in the area of the propellant
management devices 5 with a maximum permissible angle, at the end
of the last burn phase of the rocket stage. Through the
longitudinal acceleration, the liquids 6 and 7 are pressed in the
direction toward the tank bottom or floor, where the engine 4 is
arranged and where the tank outlet pipe stubs to the propellant
feed lines are located. Due to the individual angles or corners
near the tank bottom within the integrated geometry of the tank 3
and the conically shaped engine thrust frame 2, the residual
propellant quantities 6 and 7 are reduced in comparison to those in
conventional tanks. Namely, basically the total volume in the tank
interior space at the bottom of the tank has been divided between
the two chambers I and II for the oxidizer 7 and the fuel 6, so
that the residual volume of each propellant component is reduced.
Also, due to the configuration at the bottom of the two chambers I
and II, with the conical thrust frame internal part 2b arranged
inside the tank wall of the tank 3, therefore the convergence
angles at the bottom of the tank are made narrower than in a
conventional tank without the internal frame part 2b. Overall, the
integrated tank form thus leads to an approximately 50% reduction
of the residual propellant quantities 6, 7 in comparison to those
in conventional tanks. Surface tension forces, which are induced by
the existing corners, still further reduce these residual
propellant quantities, even without requiring the installation of
additional propellant management devices to achieve this.
FIGS. 3 and 4 show further enlarged illustrations of a partial area
of the arrangement according to FIG. 1, at the bottom of the tank 3
and especially in the region or chamber II with the guide plates
5a2 as well as the refillable liquid reservoir 5b2, which comprises
throughflow openings 8a covered by screens or sieves 8c, which are
supported by a disc-shaped reinforcement 8b. Note that FIG. 3 shows
only a cut-away bottom portion of the outer wall of the tank 3 and
a cut-away bottom portion of the thrust frame internal part 2b,
whereby the guide plates 5a1 have been entirely omitted and the
remainders of the tank wall 3 and internal frame part 2b have been
omitted for simplicity.
The integration of the propellant management devices 5 in the
arrangement supports the load-carrying function of the conically
shaped engine thrust frame 2 in the area of the outlets of the tank
3, for introducing the engine loads into the primary structure 1.
Namely, the propellant management devices 5 reinforce the internal
frame part 2b in the critical load introduction area at the bottom
of the tank. In that regard, the load path passes predominantly
through the conically shaped engine thrust frame internal part 2b,
so that no additional supports are needed.
Because the optimization of the diameter and of the height of a
rocket stage also depends on the configuration of the entire
launcher or booster, it is important for an optimization of the
launcher or booster capacity of a rocket stage, that the diameter
and the length of the tank can be varied. A purely spherical shape
of such a tank is not optimal in all cases, because relatively
large empty volumes can arise in the rocket stage when using a
purely spherical tank or purely hemispherical tank domes under
certain circumstances, which can cause a launcher or booster stage
to become unnecessarily large. Therefore, variants of tanks with
different shapes as illustrated in FIGS. 5 to 7 lead to better
adapted solutions in some cases. This can be a tank with a
cylindrical intermediate piece 9 between hemispherical domes as
shown in FIG. 5, a tank with essentially elliptical domes 10
according to FIG. 6, or a tank with hemispherical or partial
spherical arcuate domes 11 as illustrated in FIG. 7.
With regard to the essentially conically shaped engine thrust frame
2, the internal part 2b can be extended externally from the tank 3,
at the top by a conically shaped ring 2a (the so-called Conical
Tank Attachment Ring) for connection to the primary structure 1,
and at the bottom by a cylindrical or conical tank extension or
thrust frame extension 2c for connection to the engine 4, as shown
in FIG. 8. In that regard, the extension 2c enables a better
introduction of the actuator adjustment forces and of the engine
forces. However, this arrangement is not quite as compact and
simultaneously is more complex than one without such an extension
2c. On the other hand, in order to better utilize an existing
volume, alternatively the conically shaped ring 2a can also be
completely omitted, whereby the cylindrical primary structure 1 is
connected to the thrust frame 2 and/or to the tank 3 directly
adjoining the tank frame internal part 2b, as illustrated in FIG.
9. In such a case, however, the loads of the launcher or booster
may be partly taken over by the tank shell.
In the above described arrangements, the engine thrust frame is
embodied conically shaped with a linear conical wall. Depending on
the system requirements, however, it can also deviate from this
shape. In FIGS. 10 and 11, frames with respectively bent or curved
longitudinal surfaces of a generally conical shape are illustrated.
This can be realized with a concave curve 13 as shown in FIG. 10,
or with a convex curve as shown in FIG. 11. The selection of the
optimal bend or curvature of the longitudinal surfaces of the
engine thrust frame in that regard depends, among other things, on
the differential pressure between the two tank chambers I and II,
and further depends on the connection or junction point to the
primary structure and on the necessary volumes of the two tank
chambers I and II. In each case, in the consideration of bent or
curved longitudinal surfaces, it must be ensured that the
differential pressure is always effective in the direction of the
bend or curvature, because otherwise there is a danger of forming a
dent or bulge. When a pressure differential exists between the
chambers I and II, even an initially straight or linear walled
conical shape may bulge slightly to form a convex or concave
conical shape, which tends to equilibrate the pressure conditions.
With a curved wall shape, the load path of the engine loads is less
advantageous, because it no longer forms a straight line, and the
production of a conical structure with purposely curved walls is
also more complex.
The terms "substantially" and "essentially" when used with shapes
such as a substantially spherical shape of the tank, mean that the
shape is predominantly as defined (e.g. spherical) and it is
readily recognizable as the defined shape, but it may have slight
deviations from that shape, or it may include minor parts that
deviate from that shape.
Although the invention has been described with reference to
specific example embodiments, it will be appreciated that it is
intended to cover all modifications and equivalents within the
scope of the appended claims. It should also be understood that the
present disclosure includes all possible combinations of any
individual features recited in any of the appended claims. The
abstract of the disclosure does not define or limit the claimed
invention, but rather merely abstracts certain features disclosed
in the application.
* * * * *